This disclosure relates to a gas turbine engine. More particularly, the disclosure relates to cooling passages that may be incorporated into a gas turbine engine component, such as an airfoil, and a core used to form such cooling passages.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages formed using core structures constructed from ceramic, for example. Typically multiple cores are glued to one another and/or located within a die to provide a complex network of cooling passages.
Airfoil platforms are typically difficult to cool. Some airfoils have been proposed in which film cooling holes are provided in the platform of the blade. Such types of cooling holes may result in mixing losses and reduction in engine performance.
Still other designs utilize a separate platform core that is glued into a slot of a more substantial main core that extends from a root to an airfoil of the blade. This platform core must be relatively thin such that it can be packaged within the boundaries of the platform. As a result, the thin platform core is susceptible to breakage during the casting process, which reduces casting yield and increases part cost.